U.S. patent application number 09/735163 was filed with the patent office on 2001-10-18 for mounting for attaching a rudder to a missile.
Invention is credited to Hetzer, Walter, Lenz, Ernst.
Application Number | 20010030259 09/735163 |
Document ID | / |
Family ID | 7932920 |
Filed Date | 2001-10-18 |
United States Patent
Application |
20010030259 |
Kind Code |
A1 |
Hetzer, Walter ; et
al. |
October 18, 2001 |
Mounting for attaching a rudder to a missile
Abstract
A rudder of a missile, particularly an aircraft born ramjet
missile, is attached to the missile by a plug and socket mounting.
A bearing socket (B1) is attached to a base plate (B0) which is
secured to the missile body (FK). The socket (B1) has a conical
cavity (B3) tapering toward the base plate (B0). The plug (W1) is
formed by a rudder shaft (W) also tapering toward the base plate
and rotatably fitting into the conical cavity (B3). Bearings (L1,
L2) in the cavity (B3) hold the plug end (W1) of the rudder shaft
axially and permit rotation of the rudder shaft (W) relative to the
bearing socket (B1). The socket thus holds the bearings and has an
at least partly outer cylindrical contour with a diameter (B2) that
fits into a standard 41 mm wide slot (S) in an aircraft that
carries the missile, whereby the rudder (R) is recessed in the slot
(S) when the missile is mounted to an aircraft.
Inventors: |
Hetzer, Walter; (Grasbrunn,
DE) ; Lenz, Ernst; (Kirchseeon, DE) |
Correspondence
Address: |
FASSE PATENT ATTORNEYS, P.A.
P.O. BOX 726
HAMPDEN
ME
04444-0726
US
|
Family ID: |
7932920 |
Appl. No.: |
09/735163 |
Filed: |
December 12, 2000 |
Current U.S.
Class: |
244/3.24 ;
244/3.1 |
Current CPC
Class: |
F42B 10/64 20130101 |
Class at
Publication: |
244/3.24 ;
244/3.1 |
International
Class: |
F42B 010/00; F42B
015/01 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 16, 1999 |
DE |
199 60 738.9 |
Claims
What is claimed is:
1. An apparatus for attaching a rudder to a missile, said apparatus
comprising a mounting (B) including a base plate (B0) for securing
said rudder (R) to said missile (FK), said mounting further
comprising a bearing socket (B1) secured to said base plate (B0),
said bearing socket having a conical cavity (B3) with a small
cavity diameter next to said base plate (B0) and a large cavity
diameter at an outer end of said bearing socket (B1), said
apparatus further comprising a rudder shaft (W) for rotatably
securing said rudder through said bearing socket (B1) to said base
plate (B0), said rudder shaft (W) having a conical rudder shaft
section (W1) having a small cone diameter next to said base plate
and a large cone diameter next to said socket outer end, wherein
said conical rudder shaft section (W1) is rotatably received in
said conical cavity (B3), whereby said conical rudder shaft section
(W1) and said conical cavity (B3) taper toward said base plate
(B0), and bearings (L1, L2) mounted in said bearing socket (B1),
said bearings (L1, L2) rotatably holding said conical rudder shaft
section (W1) in said conical cavity (B3) of said socket (B1).
2. The apparatus of claim 1, further comprising a central rudder
axis (WA) extending lengthwise through said rudder shaft (W),
through said socket (B1) and through said base plate (B0), said
socket (B1) having a first bore next to said base plate, said base
plate (B0) having a second bore next to said first bore so that
said first and second bores are coaxial to said central rudder axis
(WA), a recess (B4 in said base plate (B0) communicating with said
first and second bores, said rudder shaft (W) comprising a drive
end (W5) extending through said first and second bores into said
recess (B4) in said base plate (B0), a rudder shaft drive lever (H)
rotatable in said recess and rigidly secured to said drive end (W5)
of said rudder shaft (W) for rotating said rudder (R) through said
rudder shaft (W) by rotating said drive lever (H) in said recess
(B4) of said base plate (B0).
3. The apparatus of claim 1, wherein said rudder shaft (W)
comprises a further shaft section (W2) extending out of said
conical cavity (B3) of said bearing socket (B1), said further shaft
section (W2) forming a fork with a gap (W4) for securing said
rudder (R) to said rudder shaft (W).
4. The apparatus of claim 3, wherein said further shaft section
(W2) has a configuration tapering away from said bearing socket
(B1) toward said rudder (R).
5. The apparatus of claim 3, further comprising a rudder (R)
including a rudder foot (RF), a cut-out (CO) forming rudder foot
extensions (RFE) in said rudder foot (RF), said rudder foot (RF)
extending into said gap (W4), said cut-out extending along said
bearing socket (B1) so that said rudder foot extensions (RFE) reach
toward said base plate (B0) without contacting said base plate
(B0).
6. The apparatus of claim 1, wherein said bearings are ceramic
needle bearings.
7. The apparatus of claim 1, wherein said bearing socket (B1)
comprises an outer configuration that is at least partly
cylindrical having an outer diameter (B2), said apparatus further
comprising a rudder blade with a cut-out (CO) facing said bearing
socket (B1), said outer diameter (B2) of said socket fitting into
said cut-out with a clearance between said bearing socket (B1) and
said rudder blade.
8. The apparatus of claim 1, wherein said bearings (L1, L2)
comprise a large diameter bearing (L1) mounted in said cavity (B3)
at said large cavity diameter at said outer end of said bearing
socket (B1), and a small diameter bearing (L2) mounted in said
cavity (B3) at said small cavity diameter next to said base plate
(B0).
9. The apparatus of claim 1, wherein said rudder shaft (W)
comprises a further shaft section (W2) extending out of said
conical cavity (B3) for securing said rudder (R) to said rudder
shaft (W), said rudder shaft (W) comprising a first large diameter
cylindrical neck (N1) between said conical rudder shaft section
(W1) and said further shaft section (W2), said bearings comprising
a first large diameter bearing (L1) mounted on said first large
diameter cylindrical neck (N1), said rudder shaft (W) further
comprising a second small diameter cylindrical section (W5) forming
a drive end of said rudder shaft, said bearings comprising a second
small diameter bearing (L2) mounted on said small diameter
cylindrical section (W5), said bearing socket (B1) and said base
plate (B0) comprising a bore in which said second small diameter
bearing (L2) is received.
10. The apparatus of claim 9, wherein said drive end (W5) projects
into a recess (B4) of said base plate (B0), said apparatus further
comprising a drive lever (H) rigidly connected to said drive end
(W5) of said rudder shaft (W), and wherein said drive lever (H) is
rotatable in said recess (B4).
11. The apparatus of claim 10, wherein said first and second
bearings are ceramic needle bearings.
12. The apparatus of claim 7, wherein said outer diameter (B2) of
said bearing socket (B1) is dimensioned to fit into a standard slot
(S) in an aircraft (F) carrying said missile.
Description
PRIORITY CLAIM
[0001] This application is based on and claims the priority under
35 U.S.C. .sctn.119 of German Patent Application 199 60 738.9,
filed on Dec. 16, 1999, the entire disclosure of which is
incorporated herein by reference.
FIELD OF THE INVENTION
[0002] The invention relates to a mounting for attaching a rudder
to a missile, particularly a guided missile driven by a ramjet and
carried by an aircraft. A rudder blade is mounted to a rudder shaft
which in turn is secured to an interface fitting referred to as a
"mounting" attachable to the missile body. The rudder blade is
turnable by a rudder actuating lever.
BACKGROUND INFORMATION
[0003] German Patent Publication DE 196 35 847 C2 describes a
mounting as mentioned above. Power for operating the rudder
actuating lever is transmitted from a power source through a
coupling rod with a bearing at each end of the rod. The available
space is not used efficiently and conventional thermal stress
characteristics and mechanical stress characteristics leave room
for improvement. Similar considerations apply with regard to
reloading the same aircraft with missiles, particularly different
missile types.
[0004] German Patent Publication DE 34 41 534 C2 discloses a
bearing for a rudder blade of a guided flying body that is launched
from a firing tube. The rudder is mounted in the tail end of the
flying body, whereby the mounting requires a seal for protecting
the rudder bearing against propulsion gases. The sealing pressure
is adjustable by set screws. Such a mounting is not suitable for
connecting a rudder to a missile carried by an aircraft.
[0005] Modern combat aircraft carry medium range guided missiles
mainly in a partially recessed arrangement in the fuselage to
reduce air drag and to favorably influence the radar signature of
the combat aircraft.
[0006] The shape or configuration of the airplane missile interface
where the missile or rocket is mounted to the aircraft is
determined by the currently accepted air to air guided missile
known as AMRAAM. The configuration of the AMRAAM rudder mounting
was also used in the prototypes of the EF 2000 Euro fighter
aircraft. For mounting the missile or rocket to the aircraft
slotted recesses 41 mm wide are provided in the airplane fuselage
for accepting the rudder and wings of the AMRAAM missile when the
missiles are mounted to the aircraft.
[0007] In a case of AMRAAM trailing missiles driven by a ramjet,
the rudder must be mounted outside the missile body because the
interior is almost completely taken up by the ramjet combustor or
combustion chamber. This requirement generally leads to a
voluminous mounting outside of the missile body that may be
incompatible with the aircraft interface determined by the AMRAAM
missile.
[0008] It is not sufficient to place the missile rudder contact
free in the 41 mm wide recess of the aircraft fuselage. The
required minimum free space of several millimeters between the
aircraft body and the rocket or missile must be maintained on all
sides between the rudder and the wall of the recess.
[0009] The desire for using exchangeable ramjet driven missiles on
the same aircraft interface is therefore problematic in
conventional rudder mounting configurations.
OBJECTS OF THE INVENTION
[0010] In view of the above it is the aim of the invention to
achieve the following objects singly or in combination:
[0011] to provide a rudder mounting for attaching a rudder to an
aircraft borne missile, particularly a guided missile that fulfils
the spatial requirements while simultaneously efficiently taking up
the mechanical and thermal stresses that arise at cruising speeds
of Mach IV;
[0012] to make sure that the minimal spacing between the slot walls
in the aircraft body and the missile or rocket is maintained at all
times;
[0013] to provide a rudder to rocket mounting interface that
permits quick reloading of missiles even under adverse field
conditions;
[0014] to permit quick reloading even when using exchangeable
different rockets or missiles;
[0015] to permit mounting the rocket rudder to the missile by using
standard tools and even under adverse field conditions; and
[0016] to construct a missile rudder mounting in such a way that
bending loads at the rudder connecting point of the mounting are
reduced or even minimized.
SUMMARY OF THE INVENTION
[0017] The mounting for attaching a rudder to a missile according
to the invention is characterized by an interface fitting simply
referred as "mounting" including a base plate for securing the
rudder to the missile. The mounting further comprises a bearing
socket secured to the base plate. The bearing socket has a conical
cavity with a small cavity diameter next to the base plate and a
large cavity diameter at a socket outer end facing the rudder. The
mounting further includes a rudder shaft for rotatably securing the
rudder through the socket to the base plate. The rudder shaft has a
conical rudder shaft section having a small cone diameter next to
the base plate and a large cone diameter next to the socket outer
end. The conical rudder shaft section is rotatably received in the
conical cavity, whereby the conical rudder shaft section and the
conical cavity taper toward the base plate. Bearings are mounted in
the socket and rotatably hold the conical rudder shaft section in
the conical cavity of the socket.
[0018] A radially outer end of the rudder shaft for holding a
rudder blade is provided with a fork configuration having two
prongs forming a gap in which a blade foot of the rudder is
mounted. The radially outer end of the rudder shaft formed by the
two prongs is preferably also conical.
[0019] The outer configuration of the bearing socket is at least
partially cylindrical and has such an outer diameter that it fits
with the required all around spacing into the above-mentioned slot
in an aircraft carrying the missile.
BRIEF DESCRIPTION OF THE DRAWINGS
[0020] In order that the invention may be clearly understood, it
will now be described in connection with an example embodiment,
with reference to the accompanying drawings, wherein:
[0021] FIG. 1 is a view in the direction of the longitudinal
missile axis illustrating schematically the position of a missile
rudder in a slot in an aircraft carrying the missile;
[0022] FIG. 2a is a view in the direction of the arrow A in FIG. 1
showing a plan view of the present rudder mounting with its base
plate attached to a missile body;
[0023] FIG. 2b is a sectional view along section line IIB-IIB in
FIG. 2A showing a drive lever for the rudder and the bending
moments effective on the present rudder mounting; and
[0024] FIG. 2c is a sectional view along section plane IIC-IIC in
FIG. 2A on a scale somewhat enlarged relative to FIG. 2B and
showing, broken-away, rudder blade foot extension with a cut-out in
which a bearing socket of the mounting is received.
DETAILED DESCRIPTION OF PREFERRED EXAMPLE EMBODIMENTS AND OF THE
BEST MODE OF THE INVENTION
[0025] FIG. 1 shows schematically a missile body FK mounted to an
aircraft of which only a broken-away portion F is shown. The
aircraft portion F may be part of an aircraft fuselage or a wing
and has a slot S in which a rudder R of the missile body FK is
received or enclosed. The slot S has a standardized width SW of 41
mm. The rudder R is secured to the missile body FK by a mounting B
according to the invention including a base plate BO and a bearing
socket B1 as will be described in more detail below. The rudder R
has a rudder foot RF with an outer diameter of 21 mm, thus leaving
a spacing between the rudder foot RF and the inner wall of the slot
S all around the rudder and rudder foot. The mounting B has an
axial length of 45 mm, for example. The missile body has a
diameter, for example of 180 mm and a radial distance R' from the
central axis CA of the missile body FK to the surface of the
aircraft portion F of R'=100 mm for example, thereby also leaving
the required gap between the aircraft surface and the surface of
the missile body FK. The missile is, for example of the AMRAAM
type.
[0026] To provide the required exchangeability of missiles the
exchangeable missiles must be compatible with the just described
mounting environment. More specifically, the rudder must fit in the
respective slot S with its width SW of 41 mm.
[0027] Referring to FIGS. 2a, 2b and 2c in conjunction, these Figs.
illustrate a rudder mounting B according to the invention that is
compatible with an AMRAAM missile and simultaneously useful for
securing a rudder to a missile driven by a ramjet. The present
mounting B may also be used for other missile types to provide for
the desired changeability.
[0028] FIG. 2a shows a view in which the longitudinal rudder axis
WA extends perpendicularly to the plane of the drawing sheet and
perpendicularly to the longitudinal axis CA of the missile body
FK.
[0029] The mounting B according to the invention comprises a base
plate B0 secured to the missile body FK by conventional mounting
elements B0' such as screws or rivets or the like. Only the left
portion of the rudder R is shown in full lines in FIG. 2a while the
right-hand portion R1 is shown in dashed lines. The present
mounting B further comprises a rudder shaft W partially received in
a bearing socket B1 secured to the base plate B0. The rudder shaft
has a forked end W2 with two prongs forming a gap W4 with a saddle
W3 best seen in FIG. 2b in which a rudder blade foot RF seen in
FIG. 2c, is received and secured by conventional securing elements
or connectors C such as screws, locking pins, or the like C.
[0030] The rudder blade foot RF has a cut-out CO so dimensioned
that the bearing socket B1 can be received in the cut-out CO as
best seen in FIG. 2c. For this purpose, the bearing socket B1
preferably has an at least partially cylindrical outer
configuration with a diameter B2 slightly smaller than the width of
the cut-out CO to leave sufficient clearance for rotating the
rudder R relative to the bearing socket B by operating a drive
lever H as indicated by the double arrow DA in FIG. 2a. The drive
for operating the rudder is conventional and hence not shown.
[0031] Referring particularly to FIG. 2b and 2c the bearing socket
B1 of the present mounting B has the above mentioned outer diameter
B2 and a conical cavity B3. The conical cavity B3 tapers toward the
base plate B0. The socket B1 has a bore coaxial to the rudder axis
WA. The bore forms a small diameter end of the cavity B3. The
radially outer end of the cavity B3 facing the rudder R has a large
diameter forming a cylindrical cavity section wherein a large
diameter bearing L1 is received. A small diameter bearing L2 is
received in the small diameter bore of the socket B1, whereby the
bore merges into a respective coaxial bore in the base plate B0 to
hold a small diameter bearing L2. A rudder shaft W has a conical
inner section W1 positioned in the cavity B3 and an outer section
W2 which preferably is also conical, but not necessarily. The outer
section W2 extends out of the cavity B3 to hold the rudder blade
foot RF, for example in the gap W4 formed between two tapering
sides or prongs which form an outward conical taper. The gap W4 has
a bottom forming the saddle W3 on which the rudder blade foot RF
rests.
[0032] The two rudder shaft sections W1 and W2 are interconnected
by a large diameter neck N1 having a diameter d1. The large
diameter bearing L1 is mounted on the neck N1. The opposite small
diameter end of the tapering section W1 of the rudder shaft W has a
cylindrical small diameter neck N2 with a smaller diameter d2 on
which the small diameter bearing L2 is mounted. A neck extension
projects out of the bearing L2 to form a drive end W5 of the rudder
shaft W. The projecting drive end W5 extends from the neck N2 into
a recess B4 of the base plate B0. The above mentioned drive lever H
is received in the recess B4 of the base plate B0 and is rigidly
connected to the drive end W5 of the rudder shaft W for rotating
the rudder by any conventional drive suitable for the purpose.
[0033] Referring to FIG. 2b, the just described construction of the
conical bearing socket recess B3 and the conical section W1 of the
rudder shaft W provide an advantageous distribution of the bending
moments BL and WL effective on the present mounting B. The bending
moment BL is effective on the bearing socket B1 while the bending
moment WL is effective on the rudder shaft W. These bending moments
BL and WL are caused by forces Fq exerted by the air flow on the
rudder surfaces. These forces Fq are shown as an arrow Fq in FIG.
2b. This distribution of the bending moments is very beneficial
because the larger bending moment BL is directly introduced into
the base plate B0 at the thick end of the bearing socket B1. Thus,
the dimensions of the present mounting B are effectively adapted to
taking up the bending moments BL and WL illustrated in FIG. 2b.
[0034] The bearings L1 and L2 received in the bearing socket B1 are
preferably ceramic needle bearings, at least one of which is so
constructed as to take up any possibly occurring axial forces
effective in the direction of the rudder axis WA radially
outwardly. The large diameter bearing L1 has preferably an inner
diameter of d1 of 22 mm while the smaller diameter bearing L2 has
preferably an inner diameter d2 of 12 mm, whereby the bearings are
adapted to the bending load exerted by the bending moment WL to
which the rudder shaft W is exposed.
[0035] Referring to FIGS. 2b and 2c in conjunction, the rudder R
has a rudder foot RF with a cut-out CO. The cut-out CO has an axial
length slightly smaller than the axial length of the bearing socket
B1. Further, as mentioned, the bearing socket B1 has a diameter B2
slightly smaller than the width of the cut-out CO to provide the
required clearance for the relative rotation of the rudder R
relative to the socket B1 and relative to the base plate B0. The
rudder foot RF is received in the gap W4 and rests against the
saddle W3 as mentioned to assure the axial spacing between the base
plate B0 and the inner ends of the rudder foot extensions RFE. The
rudder foot RF is rigidly secured by connectors C in the gap W4, as
mentioned above.
[0036] The radially inner end of the drive lever H is, for example,
secured in a form locking manner to the drive end W5 of the rudder
shaft W.
[0037] The rudder mounting B according to the invention has the
following advantages. The invention minimizes the dimensions of the
bearing socket B1 and the rudder shaft W while simultaneously
adapting the dimensions to the maximum bending load that can occur.
Specifically, the socket B1 has a thick dimension next to the base
plate B0 where its bending load caused by the bending moment BL is
largest. Similarly, the rudder shaft W has its largest effective
dimension where the maximum of the bending load occurs caused by
the bending moment WL.
[0038] Further, the main components such as the base plate B0, the
bearing socket B1 and the rudder shaft W are so-configured that
they can be manufactured in a cost efficient manner by simple
machining operations. This simple plug and socket configuration
also facilitates the mounting of the rudder R in the socket B1 and
the socket B1 to the base plate B0 and the base plate B0 to the
missile body FK.
[0039] The present mounting B has the further advantage that the
aerodynamic heating that occurs at speeds up to Mach IV causing a
high thermal loading, results in a homogeneous heat distribution in
the just mentioned main components of the present mounting. Such a
uniform heat distribution or uniform heating permits the
advantageous use of cost effective high temperature resistant
ceramic needle bearings L1 and L2.
[0040] Forming the outer end W2 of the rudder shaft W as a forked
configuration with the gap W4 facilitates the manufacture as well
as the rapid mounting of the rudder blade in the socket W4 even
under adverse conditions in the field, whereby standard tools can
be used.
[0041] The arrangement and position of the gap W4 in which the
rudder blade foot RF is mounted connects the rudder blade to the
rudder shaft W at the center of pressure of the rudder blade so
that the base bending moments are small at the clamping or bearing
point.
[0042] Yet another advantage is seen in the fact that a sealing
between the missile body FK and the present mounting B can be
simply arranged between the base plate B0 and the outer surface of
the missile body FK. The same applies with regard to the rudder
drive lever H, thereby obtaining a perfect sealing since the
location of the relative motion between the rudder shaft W and the
bearing socket B1 is removed from the missile body FK or rather
separated from the missile body FK by the base plate B0.
[0043] Although the invention has been described with reference to
specific example embodiments, it will be appreciated that it is
intended to cover all modifications and equivalents within the
scope of the appended claims. It should also be understood that the
present disclosure includes all possible combinations of any
individual features recited in any of the appended claims.
* * * * *